US8006477B2 - Re-heat combustor for a gas turbine engine - Google Patents
Re-heat combustor for a gas turbine engine Download PDFInfo
- Publication number
- US8006477B2 US8006477B2 US12/060,481 US6048108A US8006477B2 US 8006477 B2 US8006477 B2 US 8006477B2 US 6048108 A US6048108 A US 6048108A US 8006477 B2 US8006477 B2 US 8006477B2
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- Prior art keywords
- combustion
- turbine
- flow
- turbine engine
- engine according
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/34—Feeding into different combustion zones
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas- turbine plants for special use
- F02C6/003—Gas-turbine plants with heaters between turbine stages
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/70—Shape
- F05D2250/71—Shape curved
Definitions
- the present invention relates to the art of turbomachines and, more particularly, to a re-heat combustor for a gas turbine engine.
- gas turbine engines combust a fuel/air mixture to release heat energy to form a high temperature gas stream.
- the high temperature gas stream is channeled to a turbine section via a hot gas path.
- a compressor compresses incoming air to a high pressure.
- the high pressure air is delivered to a combustion chamber to mix with fuel and form a combustible mixture.
- the combustible mixture is then ignited to form a high pressure, high velocity gas stream that is delivered to a turbine section of the turbine engine.
- the turbine section converts thermal energy from the high temperature, high velocity gas stream to mechanical energy that rotates a turbine shaft.
- the turbine shaft is coupled to and drives the compressor and also other machinery such as an electrical generator.
- exhaust gases are formed and vented from the turbine.
- the exhaust gases can either be expelled to ambient air or used in such a way so as to recuperate a portion of energy in the exhaust gases an thus increase cycle efficiency.
- Enhancing cycle efficiency at various operating conditions, particularly at base load, and off peak loads is always a concern.
- some turbine engines employ a reheat combustor that recycles a portion of the exhaust gases in the turbine. While adding some level of efficiency, the use of a reheat combustor, particularly between turbine sections, typically increases an axial length of the turbine engine. That is, reheat combustors require additional cooling and additional flow paths for combustion. The additional flow paths result in an overall increase in turbine shaft length.
- a turbine engine constructed in accordance with an exemplary embodiment of the invention includes a turbine section having a first turbine portion and a second turbine portion arranged along a central axis.
- a re-heat combustor is arranged between the first and second turbine portions.
- the re-heat combustor includes a combustion duct having a curvilinear flow portion. The curvilinear flow portion provides an increased residence time of combustion products passing through the re-heat combustor.
- Exemplary embodiments of the present invention also include a method of operating a turbine engine.
- the method includes generating an airflow in a first turbine portion of the turbine engine, and passing the airflow toward a re-heat combustor having a combustion duct.
- the method further includes directing the air flow through a combustion flow inlet of the combustion duct, guiding the air flow along a curvilinear flow portion of the combustion duct toward a combustion flow outlet, and passing the airflow through the combustion flow outlet into a second turbine portion of the turbine engine.
- FIG. 1 is a partial, schematic representation of a turbine engine including a re-heat combustor constructed in accordance with exemplary embodiments of the present invention
- FIG. 2 is a left perspective view of the re-heat combustor of FIG. 1 ;
- FIG. 3 is a partial perspective view of a flow conduit portion of the re-heat combustor of FIG. 2 ;
- FIG. 4 is a schematic representation of a reheat combustor constructed in accordance with one exemplary embodiment of the invention.
- FIG. 5 is a schematic representation of a reheat combustor constructed in accordance with another exemplary embodiment of the invention.
- Turbine section 10 receives hot gases of combustion from an annular array of combustors (not shown), which transmit the hot gases through a transition duct or piece 12 .
- the combustion gases pass through transition duct 12 into combustion receiving zone 13 and flow toward a first or high pressure (HP) turbine portion 14 .
- HP turbine portion 14 includes a number of turbine stages one of which is indicated at 15 .
- Turbine stage 15 includes a plurality of circumferentially spaced buckets, one of which is indicated at 16 , mounted on, and forming part of, a turbine roller 18 , and a plurality of circumferentially spaced stator vanes, one of which is indicated at 20 .
- Turbine section 10 also includes a second or low pressure (LP) turbine portion 21 fluidly connected to first turbine portion 14 along a central axis (not separately labeled).
- LP turbine section 21 includes a first LP stage 22 having a plurality of circumferentially-spaced stator vanes, one of which is indicated at 23 .
- First LP stage 22 also includes a plurality of circumferentially-spaced buckets 26 mounted on a roller wheel 27 .
- LP turbine portion 21 includes a second LP stage 27 having a plurality of circumferentially spaced buckets 28 mounted on a roller wheel 30 , and a plurality of circumferentially spaced stator vanes 32 .
- the number of stages present within each turbine portion 14 and 21 can vary.
- HP turbine portion 14 is joined to LP turbine portion 21 via a re-heat combustor 42 .
- re-heat combustor 42 includes a main body portion 44 having a first end section 46 that extends to a second end section 47 through an intermediate section 48 that collectively define a plurality of flow conduits, one of which is indicated at 50 , arranged in an annular array about main body portion 44 .
- flow conduit 50 As each of the plurality of flow conduits is substantially similar, a detailed description will follow with reference to flow conduit 50 .
- Flow conduit 50 includes a bypass duct 55 and a combustion duct 58 which, in the exemplary embodiment shown, is configured in a can-annular arrangement.
- Bypass duct 55 includes a first end portion 61 that extends to a second end portion 62 through an intermediate portion 63 .
- First end portion 61 includes a bypass flow inlet 65 that is fluidly connected to HP turbine portion 14 and a bypass flow outlet 66 fluidly connected to LP turbine portion 21 .
- a first portion of a turbine air flow passes from HP turbine portion 14 , into bypass flow inlet 65 , through bypass duct 55 and exits into LP turbine portion 21 through bypass flow outlet 66 .
- combustion duct 58 includes a first end portion 70 that extends to a second end portion 71 through an intermediate portion 72 .
- Intermediate portion 72 is defined by an arcuate or curvilinear flow portion that extends outward and is spaced from bypass duct 55 .
- Combustion duct 58 further includes a combustion flow inlet 74 arranged at first end portion 70 and a combustion air outlet 75 arranged at second end portion 71 .
- combustion duct 58 is provided with a variable position flow diverter 80 arranged at combustion flow inlet 74 .
- Flow diverter 80 is pivotally mounted at first end portion 70 and selectively shifted to alter an inlet geometry of combustion flow inlet 74 in order to enhance combustion efficiency, particularly at lower turbine speeds.
- flow diverter 80 could direct all airflow through the bypass duct 55 at certain load conditions, e.g., when enthalpy addition in re-heat combustor 42 is not needed thereby reducing a pressure drop at re-heat combustor 42 and increasing gas turbine efficiency.
- combustion duct 58 is provided with a plurality of fuel injectors indicated generally at 90 that are fluidly connected to a plurality of fuel ports indicated generally at 93 (See FIG. 5 ) and at least one fuel manifold (not shown). Fuel injectors 90 are arranged proximate to first end portion 70 and are configured to introduce fuel into an air flow passing through combustion duct 58 .
- Combustion duct 58 is also provided with a plurality of dilution jets 96 arranged proximate to second end portion 71 . Dilution jets 96 fluidly connect bypass duct 55 with combustion duct 58 . Dilution jets 96 allow a portion of the air flow passing through bypass duct 55 to enter into combustion duct 58 to dilute the products of combustion before passing though combustion duct outlet 75 and into LP turbine portion 21 .
- the curvilinear flow portion of the combustion duct enhances combustion control efficiency. That is, combustion products passing though combustion duct 58 are provided with an increased residence time without expanding an axial length of re-heat combustor 42 and, by extension, minimizing an axial length of turbine portion 10 . By increasing residence time of the combustion products, re-heat combustor enhances combustion efficiency while, at the same time, maintaining emissions within compliance levels. Also, it should be recognized that the overall shape/geometry of the curvilinear flow portion can be adjusted/altered/tailored to establish a residence time for specific turbines without increasing turbine length.
- a steam jacket can be mounted about the combustion ducts to provide additional cooling and further enhance gas turbine performance (emissions, output and efficiency) as less air is utilized for cooling.
Abstract
A turbine engine includes a turbine section having a first turbine portion and a second turbine portion arranged along a central axis. A re-heat combustor is arranged between the first and second turbine portions. The re-heat combustor includes a combustion duct having a curvilinear flow portion. The curvilinear flow portion provides an increased residence time of combustion products passing through the re-heat combustor.
Description
The present invention relates to the art of turbomachines and, more particularly, to a re-heat combustor for a gas turbine engine.
In general, gas turbine engines combust a fuel/air mixture to release heat energy to form a high temperature gas stream. The high temperature gas stream is channeled to a turbine section via a hot gas path. More specifically, a compressor compresses incoming air to a high pressure. The high pressure air is delivered to a combustion chamber to mix with fuel and form a combustible mixture. The combustible mixture is then ignited to form a high pressure, high velocity gas stream that is delivered to a turbine section of the turbine engine. The turbine section converts thermal energy from the high temperature, high velocity gas stream to mechanical energy that rotates a turbine shaft. The turbine shaft is coupled to and drives the compressor and also other machinery such as an electrical generator.
After converting the thermal energy from the high pressure, high velocity gas stream to mechanical energy, exhaust gases are formed and vented from the turbine. The exhaust gases can either be expelled to ambient air or used in such a way so as to recuperate a portion of energy in the exhaust gases an thus increase cycle efficiency. Enhancing cycle efficiency at various operating conditions, particularly at base load, and off peak loads is always a concern. Towards that end, some turbine engines employ a reheat combustor that recycles a portion of the exhaust gases in the turbine. While adding some level of efficiency, the use of a reheat combustor, particularly between turbine sections, typically increases an axial length of the turbine engine. That is, reheat combustors require additional cooling and additional flow paths for combustion. The additional flow paths result in an overall increase in turbine shaft length. Extending the turbine shaft length creates efficiency losses and adds maintenance and reliability concerns. Reheat combustors also possess a high demand for cooling air. Typically, the cooling air is extracted from a compressor portion of the turbine engine. Unfortunately, as the high pressure cooling air is not used to produce work, extracting compressor air for cooling creates efficiency losses.
A turbine engine constructed in accordance with an exemplary embodiment of the invention includes a turbine section having a first turbine portion and a second turbine portion arranged along a central axis. A re-heat combustor is arranged between the first and second turbine portions. The re-heat combustor includes a combustion duct having a curvilinear flow portion. The curvilinear flow portion provides an increased residence time of combustion products passing through the re-heat combustor.
Exemplary embodiments of the present invention also include a method of operating a turbine engine. The method includes generating an airflow in a first turbine portion of the turbine engine, and passing the airflow toward a re-heat combustor having a combustion duct. The method further includes directing the air flow through a combustion flow inlet of the combustion duct, guiding the air flow along a curvilinear flow portion of the combustion duct toward a combustion flow outlet, and passing the airflow through the combustion flow outlet into a second turbine portion of the turbine engine.
Additional features and advantages are realized through the techniques of exemplary embodiments of the present invention. Other embodiments and aspects of the invention are described in detail herein and are considered a part of the claimed invention. For a better understanding of the invention with advantages and features thereof, refer to the description and to the drawings.
With initial reference to FIG. 1 , there is illustrated a representative example of a turbine section of a gas turbine engine, generally indicated at 10. Turbine section 10 receives hot gases of combustion from an annular array of combustors (not shown), which transmit the hot gases through a transition duct or piece 12. The combustion gases pass through transition duct 12 into combustion receiving zone 13 and flow toward a first or high pressure (HP) turbine portion 14. HP turbine portion 14 includes a number of turbine stages one of which is indicated at 15. Turbine stage 15 includes a plurality of circumferentially spaced buckets, one of which is indicated at 16, mounted on, and forming part of, a turbine roller 18, and a plurality of circumferentially spaced stator vanes, one of which is indicated at 20.
In accordance with the exemplary embodiment shown. HP turbine portion 14 is joined to LP turbine portion 21 via a re-heat combustor 42. As best shown in FIGS. 2-5 , re-heat combustor 42 includes a main body portion 44 having a first end section 46 that extends to a second end section 47 through an intermediate section 48 that collectively define a plurality of flow conduits, one of which is indicated at 50, arranged in an annular array about main body portion 44. As each of the plurality of flow conduits is substantially similar, a detailed description will follow with reference to flow conduit 50.
As further shown in FIGS. 2-3 combustion duct 58 includes a first end portion 70 that extends to a second end portion 71 through an intermediate portion 72. Intermediate portion 72 is defined by an arcuate or curvilinear flow portion that extends outward and is spaced from bypass duct 55. Combustion duct 58 further includes a combustion flow inlet 74 arranged at first end portion 70 and a combustion air outlet 75 arranged at second end portion 71. In addition, combustion duct 58 is provided with a variable position flow diverter 80 arranged at combustion flow inlet 74. Flow diverter 80 is pivotally mounted at first end portion 70 and selectively shifted to alter an inlet geometry of combustion flow inlet 74 in order to enhance combustion efficiency, particularly at lower turbine speeds. It should be appreciated that flow diverter 80 could direct all airflow through the bypass duct 55 at certain load conditions, e.g., when enthalpy addition in re-heat combustor 42 is not needed thereby reducing a pressure drop at re-heat combustor 42 and increasing gas turbine efficiency.
In accordance with the exemplary embodiment shown, combustion duct 58 is provided with a plurality of fuel injectors indicated generally at 90 that are fluidly connected to a plurality of fuel ports indicated generally at 93 (See FIG. 5 ) and at least one fuel manifold (not shown). Fuel injectors 90 are arranged proximate to first end portion 70 and are configured to introduce fuel into an air flow passing through combustion duct 58. Combustion duct 58 is also provided with a plurality of dilution jets 96 arranged proximate to second end portion 71. Dilution jets 96 fluidly connect bypass duct 55 with combustion duct 58. Dilution jets 96 allow a portion of the air flow passing through bypass duct 55 to enter into combustion duct 58 to dilute the products of combustion before passing though combustion duct outlet 75 and into LP turbine portion 21.
At this point it should be appreciated that the curvilinear flow portion of the combustion duct enhances combustion control efficiency. That is, combustion products passing though combustion duct 58 are provided with an increased residence time without expanding an axial length of re-heat combustor 42 and, by extension, minimizing an axial length of turbine portion 10. By increasing residence time of the combustion products, re-heat combustor enhances combustion efficiency while, at the same time, maintaining emissions within compliance levels. Also, it should be recognized that the overall shape/geometry of the curvilinear flow portion can be adjusted/altered/tailored to establish a residence time for specific turbines without increasing turbine length. In addition, by spacing the combustion duct outward from main body portion of the re-heat combustor, cooling requirements are split between the combustion duct and the bypass duct thereby allowing for increased flexibility of hot component cooling. Finally, when cooling is required, a steam jacket can be mounted about the combustion ducts to provide additional cooling and further enhance gas turbine performance (emissions, output and efficiency) as less air is utilized for cooling.
In general, this written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the exemplary embodiments of the present invention if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
Claims (13)
1. A turbine engine comprising:
a turbine section including a first turbine portion and a second turbine portion arranged along a central axis; and
a re-heat combustor arranged between the first and second turbine portions, the re-heat combustor including a combustion duct having a combustion flow inlet and a bypass flow inlet adjacent to the combustion flow inlet, the combustion flow inlet extends to a combustion flow outlet through a curvilinear flow portion having a vertex, the curvilinear flow portion extending from the combustion flow inlet to the vertex radially outwardly from the central axis, the curvilinear flow portion providing increased residence time of combustion products passing through the re-heat combustor.
2. The turbine engine according to claim 1 , wherein the curvilinear flow portion is spaced from the re-heat combustor.
3. The turbine engine according to claim 1 , wherein the combustion duct includes a combustion flow inlet is fluidly connected to the first turbine portion and the combustion flow outlet is fluidly connected to the second turbine portion.
4. The turbine engine according to claim 3 , wherein the combustion duct includes a flow diverter arranged at the combustion flow inlet.
5. The turbine engine according to claim 4 , wherein the flow diverter is pivotally mounted to the re-heat combustor, the flow diverter being selectively positioned to vary a geometry of the combustion flow inlet.
6. The turbine engine according to claim 3 , wherein the re-heat combustor includes at least one fuel injector mounted in the combustion duct.
7. The turbine engine according to claim 6 , wherein the at least one fuel injector is mounted at the combustion flow inlet.
8. The turbine engine according to claim 3 , wherein the re-heat combustor includes a bypass duct positioned adjacent the combustion duct.
9. The turbine engine according to claim 8 , wherein the re-heat combustor includes at least one dilution jet fluidly connecting the bypass duct and the combustion duct.
10. The turbine engine according to claim 9 , wherein the at least one dilution jet is mounted adjacent the combustion flow outlet.
11. The turbine engine according to claim 8 , wherein the bypass duct includes a bypass flow inlet and a bypass flow outlet, the bypass flow inlet being arranged adjacent the combustion flow inlet and the bypass flow outlet being arranged adjacent the combustion flow outlet.
12. The turbine engine according to claim 1 , wherein the re-heat combustor has one of a can-annular array and an annular geometry.
13. The turbine engine according to claim 1 , wherein the first turbine portion is a high pressure turbine portion and the second turbine portion is a low pressure turbine portion.
Priority Applications (5)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/060,481 US8006477B2 (en) | 2008-04-01 | 2008-04-01 | Re-heat combustor for a gas turbine engine |
FR0951924A FR2929338A1 (en) | 2008-04-01 | 2009-03-25 | POST-COMBUSTION CHAMBER FOR GAS TURBINE ENGINE |
JP2009075419A JP2009250605A (en) | 2008-04-01 | 2009-03-26 | Reheat combustor for gas turbine engine |
DE102009003702A DE102009003702A1 (en) | 2008-04-01 | 2009-03-30 | Reheating combustion chamber for a gas turbine |
CN200910133432.9A CN101550872A (en) | 2008-04-01 | 2009-04-01 | Re-heat combustor for a gas turbine engine |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/060,481 US8006477B2 (en) | 2008-04-01 | 2008-04-01 | Re-heat combustor for a gas turbine engine |
Publications (2)
Publication Number | Publication Date |
---|---|
US20090241505A1 US20090241505A1 (en) | 2009-10-01 |
US8006477B2 true US8006477B2 (en) | 2011-08-30 |
Family
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Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US12/060,481 Expired - Fee Related US8006477B2 (en) | 2008-04-01 | 2008-04-01 | Re-heat combustor for a gas turbine engine |
Country Status (5)
Country | Link |
---|---|
US (1) | US8006477B2 (en) |
JP (1) | JP2009250605A (en) |
CN (1) | CN101550872A (en) |
DE (1) | DE102009003702A1 (en) |
FR (1) | FR2929338A1 (en) |
Cited By (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20120180499A1 (en) * | 2011-01-14 | 2012-07-19 | General Electric Company | Power generation system |
US8256202B1 (en) * | 2008-11-25 | 2012-09-04 | Florida Turbine Technologies, Inc. | High bypass turbofan |
US9328663B2 (en) | 2013-05-30 | 2016-05-03 | General Electric Company | Gas turbine engine and method of operating thereof |
US9366184B2 (en) | 2013-06-18 | 2016-06-14 | General Electric Company | Gas turbine engine and method of operating thereof |
US9534541B2 (en) | 2013-10-11 | 2017-01-03 | General Electric Company | System and method for improving efficiency of a gas turbine engine |
US9624829B2 (en) | 2013-03-05 | 2017-04-18 | Industrial Turbine Company (Uk) Limited | Cogen heat load matching through reheat and capacity match |
US10036317B2 (en) | 2013-03-05 | 2018-07-31 | Industrial Turbine Company (Uk) Limited | Capacity control of turbine by the use of a reheat combustor in multi shaft engine |
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US8091365B2 (en) * | 2008-08-12 | 2012-01-10 | Siemens Energy, Inc. | Canted outlet for transition in a gas turbine engine |
US20110219776A1 (en) * | 2010-03-15 | 2011-09-15 | General Electric Company | Aerodynamic flame stabilizer |
US20120151935A1 (en) * | 2010-12-17 | 2012-06-21 | General Electric Company | Gas turbine engine and method of operating thereof |
RU2570480C2 (en) * | 2012-08-24 | 2015-12-10 | Альстом Текнолоджи Лтд | Mixing of diluting air in gas turbine sequential combustion system |
US11371701B1 (en) * | 2021-02-03 | 2022-06-28 | General Electric Company | Combustor for a gas turbine engine |
CN114583330B (en) * | 2022-03-15 | 2023-06-27 | 西安航空学院 | Protective box provided with efficient heat conduction cooling structure and used for automobile power battery |
Citations (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3054257A (en) * | 1953-03-10 | 1962-09-18 | Garrett Corp | Gas turbine power plant for vehicles |
US3315467A (en) * | 1965-03-11 | 1967-04-25 | Westinghouse Electric Corp | Reheat gas turbine power plant with air admission to the primary combustion zone of the reheat combustion chamber structure |
US3449914A (en) * | 1967-12-21 | 1969-06-17 | United Aircraft Corp | Variable flow turbofan engine |
US4085583A (en) * | 1975-03-31 | 1978-04-25 | The Boeing Company | Method for selectively switching motive fluid supply to an aft turbine of a multicycle engine |
US4206593A (en) * | 1977-05-23 | 1980-06-10 | Institut Francais Du Petrole | Gas turbine |
US4270342A (en) * | 1978-06-16 | 1981-06-02 | Bbc Brown, Boveri & Co. Ltd. | Method of operating a gas turbine plant |
US4592204A (en) * | 1978-10-26 | 1986-06-03 | Rice Ivan G | Compression intercooled high cycle pressure ratio gas generator for combined cycles |
US4858428A (en) * | 1986-04-24 | 1989-08-22 | Paul Marius A | Advanced integrated propulsion system with total optimized cycle for gas turbines |
US5184460A (en) * | 1991-01-30 | 1993-02-09 | The United States Of America As Represented By The Administrator, National Aeronautics And Space Administration | Multi-heat addition turbine engine |
US5557918A (en) * | 1994-06-03 | 1996-09-24 | Abb Research Ltd. | Gas turbine and method of operating it |
US6385959B1 (en) * | 1999-08-24 | 2002-05-14 | MONTOYA CéSAR AGUILERA | Gas turbine engine with increased fuel efficiency and method for accomplishing the same |
US6619026B2 (en) | 2001-08-27 | 2003-09-16 | Siemens Westinghouse Power Corporation | Reheat combustor for gas combustion turbine |
US6796130B2 (en) * | 2002-11-07 | 2004-09-28 | Siemens Westinghouse Power Corporation | Integrated combustor and nozzle for a gas turbine combustion system |
US7568335B2 (en) * | 2005-09-09 | 2009-08-04 | Alstom Technology Ltd | Gas turbogroup |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
IT1243682B (en) * | 1989-07-28 | 1994-06-21 | Gen Electric | GAS TURBO ENGINE STEAM COOLING |
-
2008
- 2008-04-01 US US12/060,481 patent/US8006477B2/en not_active Expired - Fee Related
-
2009
- 2009-03-25 FR FR0951924A patent/FR2929338A1/en active Pending
- 2009-03-26 JP JP2009075419A patent/JP2009250605A/en active Pending
- 2009-03-30 DE DE102009003702A patent/DE102009003702A1/en not_active Withdrawn
- 2009-04-01 CN CN200910133432.9A patent/CN101550872A/en active Pending
Patent Citations (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3054257A (en) * | 1953-03-10 | 1962-09-18 | Garrett Corp | Gas turbine power plant for vehicles |
US3315467A (en) * | 1965-03-11 | 1967-04-25 | Westinghouse Electric Corp | Reheat gas turbine power plant with air admission to the primary combustion zone of the reheat combustion chamber structure |
US3449914A (en) * | 1967-12-21 | 1969-06-17 | United Aircraft Corp | Variable flow turbofan engine |
US4085583A (en) * | 1975-03-31 | 1978-04-25 | The Boeing Company | Method for selectively switching motive fluid supply to an aft turbine of a multicycle engine |
US4206593A (en) * | 1977-05-23 | 1980-06-10 | Institut Francais Du Petrole | Gas turbine |
US4270342A (en) * | 1978-06-16 | 1981-06-02 | Bbc Brown, Boveri & Co. Ltd. | Method of operating a gas turbine plant |
US4592204A (en) * | 1978-10-26 | 1986-06-03 | Rice Ivan G | Compression intercooled high cycle pressure ratio gas generator for combined cycles |
US4858428A (en) * | 1986-04-24 | 1989-08-22 | Paul Marius A | Advanced integrated propulsion system with total optimized cycle for gas turbines |
US5184460A (en) * | 1991-01-30 | 1993-02-09 | The United States Of America As Represented By The Administrator, National Aeronautics And Space Administration | Multi-heat addition turbine engine |
US5557918A (en) * | 1994-06-03 | 1996-09-24 | Abb Research Ltd. | Gas turbine and method of operating it |
US6385959B1 (en) * | 1999-08-24 | 2002-05-14 | MONTOYA CéSAR AGUILERA | Gas turbine engine with increased fuel efficiency and method for accomplishing the same |
US6619026B2 (en) | 2001-08-27 | 2003-09-16 | Siemens Westinghouse Power Corporation | Reheat combustor for gas combustion turbine |
US6796130B2 (en) * | 2002-11-07 | 2004-09-28 | Siemens Westinghouse Power Corporation | Integrated combustor and nozzle for a gas turbine combustion system |
US7568335B2 (en) * | 2005-09-09 | 2009-08-04 | Alstom Technology Ltd | Gas turbogroup |
Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8256202B1 (en) * | 2008-11-25 | 2012-09-04 | Florida Turbine Technologies, Inc. | High bypass turbofan |
US20120180499A1 (en) * | 2011-01-14 | 2012-07-19 | General Electric Company | Power generation system |
US8322141B2 (en) * | 2011-01-14 | 2012-12-04 | General Electric Company | Power generation system including afirst turbine stage structurally incorporating a combustor |
US9624829B2 (en) | 2013-03-05 | 2017-04-18 | Industrial Turbine Company (Uk) Limited | Cogen heat load matching through reheat and capacity match |
US10036317B2 (en) | 2013-03-05 | 2018-07-31 | Industrial Turbine Company (Uk) Limited | Capacity control of turbine by the use of a reheat combustor in multi shaft engine |
US9328663B2 (en) | 2013-05-30 | 2016-05-03 | General Electric Company | Gas turbine engine and method of operating thereof |
US9366184B2 (en) | 2013-06-18 | 2016-06-14 | General Electric Company | Gas turbine engine and method of operating thereof |
US9534541B2 (en) | 2013-10-11 | 2017-01-03 | General Electric Company | System and method for improving efficiency of a gas turbine engine |
Also Published As
Publication number | Publication date |
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DE102009003702A1 (en) | 2009-10-08 |
US20090241505A1 (en) | 2009-10-01 |
CN101550872A (en) | 2009-10-07 |
JP2009250605A (en) | 2009-10-29 |
FR2929338A1 (en) | 2009-10-02 |
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